Current state of the art spacecraft fabrication and assembly techniques use a high degree of composites. Composites offer many advantages over the metallic materials commonly used in the primary supporting structure of conventional spacecraft designs. A key benefit of composites is that they can provide significant weight reductions in the final product. Composites are more easily fashioned into structural pieces of complex geometry (e.g. rounded surfaces. irregular profiles, etc.), and thus extensive use of composites can help to reduce part count as well as reduce the number of mechanical fasteners that are required to secure the parts or structural pieces together.
Examples of state of the art modular spacecraft that utilize a high degree of composites include the A2100 satellite and the IRIDIUM.RTM. Bus, both designed and built by the Lockheed Martin Corporation. In these known satellite designs, the major structural components are assembled from composite parts and bolted together, including a series of access panels. In order to reduce weight, composites are used for the spacecraft.
Although such state of the art spacecraft designs have made great strides in reducing part count and weight, the use of composites in these spacecraft designs still relies heavily on geometry and joint methods from traditional metal fabrication and assembly techniques. The result is that these spacecraft designs do not take full advantage of the benefits of working with composites. It is well understood that the efficiency of composites is decreased as the number of fasteners and discontinuous joints are increased in the completed structure.
To obtain the maximum benefits of composites, spacecraft designers must rethink the way they fabricate and assemble the major structural components of spacecraft. A continuous primary structure having a minimum number of bolt together fasteners would be a more efficient use of composites. Of course, a continuous composite containment structure for satellite primary structure fabrication is of little benefit if it does not permit good access to the spacecraft interior for installation and testing of the spacecraft subsystems prior to launch.
Inadequate or limited access to the spacecraft interior is a problem associated with most state of the art modular spacecraft designs. A case in point is the installation procedure that is required for installing the propulsion subsystem in the A2100 and IRIDIUM.RTM. spacecraft. In view of the limited access to the spacecraft interior, and further in view of the presence of other pre-installed subsystems, the propulsion subsystem must be installed as a number of subassemblies, each of size small enough to fit within the access panel openings. Once the propulsion system subassemblies are inside the access panels, they must be maneuvered around the other pre-installed subsystems into their assigned locations. The various subassemblies of the propulsion subsystem are then welded together in place. Field welding of this nature is both costly and time intensive since it must be done in a clean room environment and it further requires use of special portable welding apparatus so as not to compromise the other subsystems of the spacecraft.
Much greater manufacturing and assembling efficiencies could be realized if the propulsion subsystem could be installed within the spacecraft as a fully assembled unit.
Accordingly, a modular spacecraft fabrication and assembling method that reduces part count and number of fasteners that can also provide a structural geometry that enables full and unimpeded access to the various subsystems of the spacecraft during assembly, installation and testing would constitute a significant advancement.
The major subsystems of current spacecraft requires substantial electronics. In accordance with the conventional practice, a dedicated box of electronics is provided for each subsystem (e.g. guidance, navigation and control, command, data handling, power, propulsion, communications, thermal, etc.). The dedicated boxes are connected to each other by a series of cables. In view of the limited amount of space available within the spacecraft, the boxes are not always optimally located near their dedicated subsystems.
Further, there is significant electronic duplication associated with the dedicated box approach as each dedicated box requires a certain amount of electronics common to all dedicated boxes, such as electronics for general housekeeping functions, power supply and the like. In view of the many dedicated boxes used in a typical spacecraft and the amount of cabling required to connect all the various dedicated boxes, there are significant weight and volume penalties associated with the dedicated box approach. Also, the extra cabling introduces inefficiencies in power and signal strength.
It would be desirable to redistribute the subsystem electronics within the spacecraft in a way that reduces duplication of function and that also reduces the amount of cabling required so that a greater weight savings could be realized and power and signal loss could be minimized and thermal dissipation efficiency can be maximized.
As is well known from the PCI architecture prevalent in personal computers, electronic components can be designed as integrated circuits on standardized interface cards which, in turn, plug into a common backplane or motherboard. The improvements in performance and diagnostic advantages associated with the plug and play approach of the PCI architecture are well understood.
It would be desirable to incorporate a similar plug and play philosophy to the arrangement of the subsystem electronics on a spacecraft. More particularly, it would be desirable to arrange the electronics of the various subsystems of the spacecraft in a distributed fashion on a common backplane in a way which maximizes power, thermal and weight efficiencies.
In satellite development, a significant portion of the total cost is due to electronics qualification. The electronics must be qualified to withstand all the environmental factors associated with the space launch, such as vibration, acoustics, flame, smoke, acceleration forces, etc. In addition, the electronics must be qualified to withstand the environmental factors associated with operation in space, namely, radiation and electromagnetic interference (EMI). On top of all this, the electronics of a spacecraft must be qualified for reliable long term operation (e.g. 10-15 years) without failure.
Accordingly, a distributed electronics architecture for a spacecraft that is easily and economically space-qualified would constitute another significant advance in the art.